Endothermic fluid based thermal management system

ABSTRACT

A thermal management system for a propulsion engine of a supersonic or hypersonic aircraft is provided that uses a single flow of endothermic fluid as engine fuel and as a heat sink for engine cooling. The system includes a plurality of heat exchangers in flow series arrangement. Each heat exchanger has a reactor portion, containing a catalyst, in heat exchange relation with a heat source portion. The single fluid flow flows through each of the reactor portions and heat source portions. The heat for the reaction in the reactor portions is provided by the fluid in the heat source portion which as a result is cooled. This cooled fluid is then reheated by flowing it past a hot portion of the engine after which the fluid can flow to another reactor portion or to the engine&#39;s combustor where it is ignited.

TECHNICAL FIELD

This invention relates to a thermal energy management system for coolingengine components of supersonic and hypersonic aircraft in which asingle flow of an endothermic fluid is used as the heat sink as well asfuel for the engine.

BACKGROUND OF THE INVENTION

Thermal management systems are used to cool components, such as turbineblades, in aircraft propulsion engines. For aircraft operating at Mach 3or less these systems employ bleed air from the engine as the heat sink.As the aircraft exceeds Mach 3, the air stagnation temperature, in theengine, rises rapidly rendering the bleed air ineffective as a heatsink. An obvious alternate heat sink is the engine's fuel. However, therelatively low thermal stability temperature of conventional jet enginefuels limits their heat sink capability to temperatures at which theybegin to coke. This coking temperature is about 450° K. (350° F.) forJet Propellant(JP)-4 and about 650° K. (700° F.) for JP-7. In aircraftoperating at speeds greater than Mach 3the temperature of the engine'scomponents can easily exceed these limits. In such cases the fuel cannotbe used directly for cooling but instead must be used to cool bleed airor some other stable fluid in an intermediate cooling loop. Theintermediate cooling loop requires, for safety reasons, buffered heatexchangers which have a buffer space between the fuel and the bleed airwhich needs to be monitored for leaks of either fuel or oxidant.

Endothermic fuels are fluids which with the addition of heat and in thepresence of a catalyst decompose into two, or more, gaseous compounds atleast one of which is combustible. Because of their sensible and latentheat capacities and their endothermic chemical reaction energyabsorbtion mechanism,endothermic fuels have higher heat sinkcapabilities than Jet Propellants.

Accordingly, for aircraft operating at speeds greater than Mach 3thereis a need for an engine thermal management system that uses anendothermic fluid as a heat sink and as fuel for the engine.

SUMMARY OF THE INVENTION

An objective of the present invention is to provide an engine thermalmanagement system for a propulsion engine mounted in an aircraft capableof supersonic and hypersonic flight.

Another objective of the present invention is to provide an enginethermal management system that uses a single flow of an endothermicfluid as engine fuel and as a heat sink for engine cooling.

Yet another objective of the present invention is to provide an enginethermal management system that does not require intermediate coolingloops.

Yet still another objective of the present invention is to provide amethod in which a single flow of an endothermic fluid is used to fueland cool a propulsion engine.

The present invention achieves these objectives by providing a thermalmanagement system for a propulsion engine that uses a single flow of anendothermic fluid as engine fuel and as a heat sink for engine cooling.The system includes a plurality of heat exchangers in flow seriesarrangement. Each heat exchanger has a reactor portion containing acatalyst in heat exchange relation with a heat source portion. Thesingle fluid flow flows through each of the reactor portions and heatsource portions. The heat for the reaction in the reactor portions isprovided by the fluid in the heat source portion which as a result iscooled. This cooled fluid is then reheated by flowing it past a hotportion of the engine. After exiting the last heat exchanger, fluid isdelivered to the engine's combustor where it is ignited.

These and other objects, features and advantages of the presentinvention are specifically set forth in or Will become apparent from thefollowing detailed description of a preferred embodiment of theinvention when read in conjunction with the accompanying drawing.

BRIEF DESCRIPTION OF THE DRAWINGS

The sole figure is a schematic of the thermal management systemcontemplated by the present invention in conjunction with a propulsionengine.

DESCRIPTION OF THE PREFERRED EMBODIMENT

Referring to the drawing, a thermal management system generally denotedby reference numeral 10 is shown in conjunction with a propulsion engine12. For illustrative purposes, the propulsion engine 10 is shown as ascramjet. However, the configuration of the engine 12 can be any one ofa plurality of engine configurations such as, but not limited to, aturbojet, a ramjet, a turboramjet, or a combined cycle engine. Thepropulsion engine 12 is comprised of, in a flow series arrangement, adiffuser 14, a combustor 16, and a nozzle 18.

The thermal management system 10 includes a fuel tank 20 having anendothermic fuel, as a low pressure liquid, stored therein. A pump 22pressurizes the fuel to a sufficient pressure to satisfy the pressurerequirement of the fuel injectors, (not shown), in the combustor 16after taking into account all the pressure losses in thermal managementsystem 10. Though not necessary, to avoid problems associated with2-phase flow, this pressure should be above the critical pressure of thefuel. The fuel is now supercritical and flows through a conduit 24 to apreheater 26 in which its temperature is raised to a level at which anendothermic reaction can occur. The particular temperature level dependson the particular fuel, catalyst, and pressure. The drawing depicts thepreheater 26 as a component of the compressor 14, however the preheater26 can be any relatively low temperature heat source such as avionics,the cockpit's environmental control system, or a hollow structural panelon the engine 12. From the preheater 26, the liquid fuel flows throughconduit 28 to a first catalytic heat exchanger 30 preferably mountedaway from the engine 12.

The heat exchanger 30 has a reactor 32 in heat exchange relation with aheat source 34. The heat exchanger 30 can be a tubular type heatexchanger, or a plate-fin type heat exchanger. In a manner familiar tothose skilled in the art, the catalyst can be coated onto a surface, orplaced as packed beds in the tubes or plate fin passages of the reactor32. There being no oxidant present there is no need for buffered heattransfer. The reactor 32 receives the fuel from the conduit 28 anddelivers the fuel through conduit 36 to a second engine component 38that also requires cooling. From the component 38 the fuel returnsthrough conduit 40 to the heat source 34. At this point the fuel exitingthe heat source 34 could be sent to another component of the engine,such as 58, and then to the combustor 16. To achieve the full use of thepotential heat sink of the endothermic fuel, it is preferred to add asecond catalytic heat exchanger 44, also mounted away from the engine12, and in flow series with the heat exchanger 30.

The heat exchanger 44 has a reactor 46 in heat exchange relation with aheat source 48. A conduit 42 brings the fuel from the heat source 34 tothe reactor 46. After leaving the reactor 46, the fuel passes throughconduit 50 to another engine component 52 and then through conduit 54 tothe heat source 48. A conduit 56 transports the fuel from the heatsource 48 to an engine component 58. From the component 58 the fuelflows through a conduit 60 to the combustor 16 where it is ignited.Alternatively, additional catalytic heat exchanger reactors can be addedto the thermal management system 10.

Endothermic fuels are combustible fluids which in the presence of acatalyst and heat decompose into two, or more, gaseous compounds, atleast one of which is combustible. The endothermic fuel may be selectedfrom a group of hydrocarbon fuels which includes methylcyclohexane(MCH), decalin (hydrogenated naphthalene), methanol, n-heptane, and JetPropellant(JP) 10. Depending on the fuel selection, the catalyst isselected from a group including platinum, nickel, and zeolite(s).However, the particular selection of fuel and catalyst should not beconsidered limiting of the invention or the appended claims.

The preferred fuel and catalyst are MCH and platinum. In the presence ofheat and platinum, MCH reforms into hydrogen and tolulene, both of whichare combustible.

    C.sub.7 H.sub.14 ------ >3H.sub.2 +C.sub.7 H.sub.8         (1)

In operation, liquid MCH is pumped from the tank 20 to a pressure ofabout 4.14 MPa (600 psia) which is typically greater than its criticalpressure of 3.48 MPa (504 psia). The MCH is now a super critical fluid.Traversing the preheater 26 the fuel adsorbs about 1160 kJ/kg (500BTU/lbm) of energy raising its temperature a moderate amount to about665° K. (740° F.). As the fuel passes through the reactor 32 it adsorbsanother 1395 kJ/kg (600 BTU/lbm) from the heat source 34 and reacts withthe catalyst so that about 40% of the MCH is reformed into tolulene andhydrogen. The fuel and reaction products exit the reactor 32 at atemperature of about 835° K. (1940° F.).

The reaction in the reactor 32 occurs in a relatively narrow temperatureband due to the requirements of chemical equilibrium and kinetics,(favored by high temperatures), and catalyst life, (favored by lowtemperatures). The pressure of the fuel in the reactor 32 is also afactor as low pressures favor the endothermic reaction while highpressures hinder the formation of coke precursors. The selection of thebest pressure and temperatures for the reactor depend on the particularfuel and the requirements of the system 10. Therefore, the numberspresented here for MCH are illustrative only.

After leaving the reactor 32, the MCH traverses the engine component 38adsorbing about 815 kJ/kg (350 BTU/lbm) and exiting at a temperature ofabout 1055° K. (1440° F.). The presence of hydrogen in the partiallyreformed MCH, and the use of passivated materials in the component 38minimizes fuel stability problems. Returning to the heat source 34, 1395kJ/kg (600 BTU/lbm) is extracted and the fuel is cooled back down to1055° K. (740° F.). The amount of heat transferred from the heat source34 to the reactor 32 is limited by temperature limits of the fuel whenin contact with the engine components.

The fuel now proceeds to the reactor 46 where the fuel is again reheatedto 835° K. (1040° F.) by adsorbing 1395 kJ/kg (600 BTU/lbm) andreforming an additional 40% so that now 80% of the fuel is hydrogen andtolulene. The remaining steps being identical to those that occurredwithin the heat exchanger 30, until the fuel exits the heat source 48 ata temperature of 1055° K. (740° F.). From here the fuel passes throughthe engine component 58 where is is reheated and then sent to thecombustor 16 where it is ignited.

Thus a thermal management system for supersonic and hypersonic aircraftis provided that uses an endothermic fuel as a heat sink. By using theendothermic fuel, cooling can be accomplished by direct contact of thefuel and the engine components. Further, as the same flow of fuel isused on both sides of the heat exchangers there is no need for gaseousintermediate loops, liquid intermediate loops, or buffered heatexchangers.

Various modifications and alterations to the above described system willbe apparent to those skilled in the art. Accordingly, the foregoingdetailed description of the preferred embodiment of the invention shouldbe considered exemplary in nature and not as limiting to the scope andspirit of the invention as set forth in the following claims.

What is claimed is:
 1. A thermal energy management system for apropulsion engine mounted on an aircraft operable at high Mach numbers,said aircraft having a source of endothermic fuel in a liquid state, andsaid engine having combustor for combusting said endothermic fuel,comprising:a preheater receiving a flow of said pressurized endothermicfuel from said source; a first reactor receiving said flow from saidpreheater; a first reheater receiving said flow from said first reactor;and a first heat source receiving said flow from said first reheater,and delivering said flow to said combustor, said first heat source beingin heat exchange relation with said first reactor.
 2. The thermal energymanagement of claim 1 further including a final reheater disposedbetween said first heat source and said combustor.
 3. The thermal energymanagement of claim 2 further including a pump disposed between saidendothermic fuel source and said preheater.
 4. The thermal energymanagement system of claim 3 further comprising:a second reactorreceiving said flow from said first heat source; a second reheaterreceiving said flow from said second reactor; and a second heat sourcereceiving said flow from said second reheater, and delivering said flowto said final reheater, said second heat source being in heat exchangerelation with said second reactor.
 5. The thermal energy managementsystem of claim 4 wherein said preheater and said reheaters arestructural members of said engine requiring cooling.
 6. The thermalenergy management system of claim 4 wherein said reactors and said heatsources are mounted in said aircraft away from said engine.